U.S. patent application number 15/107493 was filed with the patent office on 2016-11-10 for fan cooling hole array.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is UNITED TECHNOLOGIES CORPORATION. Invention is credited to Matthew A. Devore, Steven Bruce Gautschi, Dominic J. Mongillo, Lane Thornton.
Application Number | 20160326883 15/107493 |
Document ID | / |
Family ID | 54333392 |
Filed Date | 2016-11-10 |
United States Patent
Application |
20160326883 |
Kind Code |
A1 |
Thornton; Lane ; et
al. |
November 10, 2016 |
FAN COOLING HOLE ARRAY
Abstract
A gas turbine engine component comprises an airfoil with a
suction side and pressure side extending from a leading edge to a
trailing edge. There are a plurality of cooling holes adjacent the
leading edge, with the cooling holes having a non-circular shape,
with a longer dimension and a smaller dimension. The airfoil
defines a radial direction from a radially outer end to a radially
inner end, and radially outer of the cooling holes spaced toward
the radially outer end, which have the longer dimension extending
closer to parallel to the radial direction. Radially inner cooling
holes closer to the radially inner end having the longer dimension
extend to be closer to perpendicular relative to the radial
direction compared to the radially outer cooling holes.
Inventors: |
Thornton; Lane; (Meriden,
CT) ; Devore; Matthew A.; (Cromwell, CT) ;
Mongillo; Dominic J.; (West Hartford, CT) ; Gautschi;
Steven Bruce; (Naugatuck, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
UNITED TECHNOLOGIES CORPORATION |
Farmington |
CT |
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Farmington
CT
|
Family ID: |
54333392 |
Appl. No.: |
15/107493 |
Filed: |
January 2, 2015 |
PCT Filed: |
January 2, 2015 |
PCT NO: |
PCT/US2015/010010 |
371 Date: |
June 23, 2016 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61928105 |
Jan 16, 2014 |
|
|
|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
Y02T 50/676 20130101;
F05D 2220/32 20130101; F05D 2260/202 20130101; F05D 2250/312
20130101; F05D 2250/313 20130101; Y02T 50/673 20130101; F01D 5/186
20130101; Y02T 50/60 20130101; F05D 2250/38 20130101; F05D 2250/14
20130101; F05D 2250/314 20130101; F05D 2250/324 20130101; F05D
2240/303 20130101 |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A gas turbine engine component comprising: an airfoil with a
suction side and pressure side, and extending from a leading edge
to a trailing edge; and a plurality of cooling holes adjacent the
leading edge, with said cooling holes having a non-circular shape,
with a longer dimension and a smaller dimension, and said airfoil
defining a radial direction from a radially outer end to a radially
inner end, and radially outer of said cooling holes spaced toward
said radially outer end having said longer dimension extending
closer to parallel to said radial direction, and radially inner
cooling holes closer to said radially inner end having said longer
dimension extending to be closer to perpendicular relative to said
radial direction compared to said radially outer cooling holes.
2. The gas turbine engine component as set forth in claim 1,
wherein said component has a platform which defines said radially
inner end.
3. The gas turbine engine component as set forth in claim 2,
wherein there is a transition zone intermediate said radially inner
and radially outer cooling holes with said transition zone
including a cooling hole having a longer direction that is
non-perpendicular and non-parallel to said radial dimension.
4. The gas turbine engine component as set forth in claim 3,
wherein said transition zone includes a plurality of cooling holes
which have said longer dimension defining an angle between 0 and
90.degree. relative to said radial dimension.
5. The gas turbine engine component as set forth in claim 4,
wherein the angle of said plurality of cooling holes in said
transition zone increasing as said cooling holes are radially
closer to said radially inner most cooling hole.
6. The gas turbine engine component as set forth in claim 5,
wherein said component is a turbine blade.
7. The gas turbine engine component as set forth in claim 6,
wherein said airfoil is a high lift airfoil.
8. The gas turbine engine component as set forth in claim 7,
wherein said high lift airfoil has a ratio of static pressure to
total pressure in proximity to an airfoil surface that is greater
than approximately 0.9 across a substantial portion of the airfoil
surface.
9. The gas turbine engine component as set forth in claim 1,
wherein there is a transition zone intermediate said radially inner
and radially outer cooling holes with said transition zone
including a cooling hole having a longer direction that is
non-perpendicular and non-parallel to said radial dimension.
10. The gas turbine engine component as set forth in claim 9,
wherein said transition zone includes a plurality of cooling holes
which have said longer dimension defining an angle between 0 and
90.degree. relative to said radial dimension.
11. The gas turbine engine component as set forth in claim 10,
wherein the angle of said plurality of cooling holes in said
transition zone increasing as said cooling holes move radially
closer to said radially inner most cooling hole.
12. The gas turbine engine component as set forth in claim 11,
wherein said component is a turbine blade.
13. The gas turbine engine component as set forth in claim 12,
wherein said airfoil is a high lift airfoil.
14. The gas turbine engine component as set forth in claim 13,
wherein said high lift airfoil has a ratio of static pressure to
total pressure in proximity to an airfoil surface that is greater
than approximately 0.9 across a substantial portion of the airfoil
surface.
15. The gas turbine engine component as set forth in claim 1,
wherein said component is a turbine blade.
16. The gas turbine engine component as set forth in claim 15,
wherein said airfoil is a high lift airfoil.
17. The gas turbine engine component as set forth in claim 16,
wherein said high lift airfoil has a ratio of static pressure to
total pressure in proximity to an airfoil surface that is greater
than approximately 0.9 across a substantial portion of the airfoil
surface.
18. The gas turbine engine component as set forth in claim 1,
wherein said airfoil is a high lift airfoil.
19. The gas turbine engine component as set forth in claim 19,
wherein said high lift airfoil has a ratio of static pressure to
total pressure in proximity to an airfoil surface that is greater
than approximately 0.9 across a substantial portion of the airfoil
surface.
20. A gas turbine engine comprising: a turbine and a compressor,
said turbine including blades and vanes; at least one of said
blades and said vanes including an airfoil with a suction side and
pressure side, and extending from a leading edge to a trailing
edge; and a plurality of cooling holes adjacent the leading edge,
with said cooling holes having a non-circular shape, with a longer
dimension and a smaller dimension, and said airfoil defining a
radial direction from a radially outer end to a radially inner end;
and there being a transition zone between said radially outer end
and said radially inner end, with said transition zones including a
plurality of cooling holes having said longer direction being
non-perpendicular and non-parallel to said radial dimension, and
said longer dimension for each of said plurality of cooling holes
defining an angle relative to said radial dimension, with the angle
of said plurality of cooling holes in said transition zone
increasing as said cooling holes move radially closer to said
radially inner end.
Description
CROSS-REFERENCE TO RELATED APPLICATION
[0001] This application claims priority to U.S. Provisional Patent
Application No. 61/928,105, filed Jan. 16, 2014.
BACKGROUND OF THE INVENTION
[0002] This application relates to a cooling hole array for use
adjacent a leading edge of an airfoil.
[0003] Gas turbine engines are known and, typically, include a fan
delivering air into a compressor where it is compressed. The
compressed air is delivered into a combustion section where it is
mixed with fuel and ignited. Products of this combustion pass
downstream over turbine rotors driving them to rotate.
[0004] The turbine section typically includes rotating blades and
static vanes, all of which include airfoils. The airfoils are
exposed to very hot temperatures and, thus, internal cooling
passages are provided within the airfoils.
[0005] Airfoils extend from a leading edge to a trailing edge and
have a suction and pressure side. Cooling has typically been
provided adjacent the leading edge with so-called "showerhead" hole
shapes. A showerhead shape has a longer dimension and a shorter
dimension and the longer dimension is typically aligned with a
radial direction along the airfoil.
[0006] More recently, so-called "high lift" aerodynamic airfoils
have been developed. A high lift airfoil maximizes the ability of
the turbine to capture energy from the working fluid, thereby
reducing the need for increasing the flow of working fluid through
the turbine and increasing turbine performance.
[0007] The use of showerhead cooling holes at the leading edge of
high lift airfoils requires some unique characteristics at least
partially because the cooling holes have typically been provided at
a mechanical leading edge, whereas the actual aerodynamics in a
high lift airfoil result in a distinct location of airflow split
extending onto the pressure side of the airfoil and, in particular,
at radially inner locations.
SUMMARY OF THE INVENTION
[0008] In a featured embodiment, a gas turbine engine component
comprises an airfoil with a suction side and pressure side
extending from a leading edge to a trailing edge. There are a
plurality of cooling holes adjacent the leading edge, with the
cooling holes having a non-circular shape, with a longer dimension
and a smaller dimension. The airfoil defines a radial direction
from a radially outer end to a radially inner end, and radially
outer of the cooling holes spaced toward the radially outer end,
which have the longer dimension extending closer to parallel to the
radial direction. Radially inner cooling holes closer to the
radially inner end having the longer dimension extend to be closer
to perpendicular relative to the radial direction compared to the
radially outer cooling holes.
[0009] In another embodiment according to the previous embodiment,
the component has a platform which defines the radially inner
end.
[0010] In another embodiment according to any of the previous
embodiments, there is a transition zone intermediate the radially
inner and radially outer cooling holes. The transition zone
includes a cooling hole having a longer direction that is
non-perpendicular and non-parallel to the radial dimension.
[0011] In another embodiment according to any of the previous
embodiments, the transition zone includes a plurality of cooling
holes which have the longer dimension defining an angle between 0
and 90.degree. relative to the radial dimension.
[0012] In another embodiment according to any of the previous
embodiments, the angle of the plurality of cooling holes in the
transition zone increase as the cooling holes are radially closer
to the radially inner most cooling hole.
[0013] In another embodiment according to any of the previous
embodiments, the component is a turbine blade.
[0014] In another embodiment according to any of the previous
embodiments, the airfoil is a high lift airfoil.
[0015] In another embodiment according to any of the previous
embodiments, the high lift airfoil has a ratio of static pressure
to total pressure in proximity to an airfoil surface that is
greater than approximately 0.9 across a substantial portion of the
airfoil surface.
[0016] In another embodiment according to any of the previous
embodiments, there is a transition zone intermediate the radially
inner and radially outer cooling holes. The transition zone
includes a cooling hole having a longer direction that is
non-perpendicular and non-parallel to the radial dimension.
[0017] In another embodiment according to any of the previous
embodiments, the transition zone includes a plurality of cooling
holes which have the longer dimension that define an angle between
0 and 90.degree. relative to the radial dimension.
[0018] In another embodiment according to any of the previous
embodiments, the angle of the plurality of cooling holes in the
transition zone increases as the cooling holes move radially closer
to the radially inner most cooling hole.
[0019] In another embodiment according to any of the previous
embodiments, the component is a turbine blade.
[0020] In another embodiment according to any of the previous
embodiments, the airfoil is a high lift airfoil.
[0021] In another embodiment according to any of the previous
embodiments, the high lift airfoil has a ratio of static pressure
to total pressure in proximity to an airfoil surface that is
greater than approximately 0.9 across a substantial portion of the
airfoil surface.
[0022] In another embodiment according to any of the previous
embodiments, the component is a turbine blade.
[0023] In another embodiment according to any of the previous
embodiments, the airfoil is a high lift airfoil.
[0024] In another embodiment according to any of the previous
embodiments, the high lift airfoil has a ratio of static pressure
to total pressure in proximity to an airfoil surface that is
greater than approximately 0.9 across a substantial portion of the
airfoil surface.
[0025] In another embodiment according to any of the previous
embodiments, the airfoil is a high lift airfoil.
[0026] In another embodiment according to any of the previous
embodiments, the high lift airfoil has a ratio of static pressure
to total pressure in proximity to an airfoil surface that is
greater than approximately 0.9 across a substantial portion of the
airfoil surface.
[0027] In another embodiment according to any of the previous
embodiments, a gas turbine engine comprises a turbine and a
compressor, the turbine including blades and vanes. At least one of
the blades and the vanes includes an airfoil with a suction side
and pressure side, and extends from a leading edge to a trailing
edge. A plurality of cooling holes are adjacent the leading edge.
The cooling holes have a non-circular shape, with a longer
dimension and a smaller dimension. The airfoil defines a radial
direction from a radially outer end to a radially inner end. There
is a transition zone between the radially outer end and the
radially inner end. The transition zones include a plurality of
cooling holes having the longer direction being non-perpendicular
and non-parallel to the radial dimension, and the longer dimension
for each of the plurality of cooling holes defining an angle
relative to the radial dimension. The angle of the plurality of
cooling holes in the transition zone increases as the cooling holes
move radially closer to the radially inner end.
[0028] These and other features may be best understood from the
following drawings and specification.
BRIEF DESCRIPTION OF THE DRAWINGS
[0029] FIG. 1 schematically shows a gas turbine engine.
[0030] FIG. 2 shows a prior art blade.
[0031] FIG. 3 shows an inventive blade.
[0032] FIG. 4A shows a feature of the prior art.
[0033] FIG. 4B schematically shows a detail of the inventive
blade.
DETAILED DESCRIPTION
[0034] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flow path B in a bypass duct defined within a
nacelle 15, while the compressor section 24 drives air along a core
flow path C for compression and communication into the combustor
section 26 then expansion through the turbine section 28. Although
depicted as a two-spool turbofan gas turbine engine in the
disclosed non-limiting embodiment, it should be understood that the
concepts described herein are not limited to use with two-spool
turbofans as the teachings may be applied to other types of turbine
engines including three-spool architectures.
[0035] The exemplary engine 20 generally includes a low speed spool
30 and a high speed spool 32 mounted for rotation about an engine
central longitudinal axis A relative to an engine static structure
36 via several bearing systems 38. It should be understood that
various bearing systems 38 at various locations may alternatively
or additionally be provided, and the location of bearing systems 38
may be varied as appropriate to the application.
[0036] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a first (or low) pressure compressor
44 and a first (or low) pressure turbine 46. The inner shaft 40 is
connected to the fan 42 through a speed change mechanism, which in
exemplary gas turbine engine 20 is illustrated as a geared
architecture 48 to drive the fan 42 at a lower speed than the low
speed spool 30. The high speed spool 32 includes an outer shaft 50
that interconnects a second (or high) pressure compressor 52 and a
second (or high) pressure turbine 54. A combustor 56 is arranged in
exemplary gas turbine 20 between the high pressure compressor 52
and the high pressure turbine 54. A mid-turbine frame 57 of the
engine static structure 36 is arranged generally between the high
pressure turbine 54 and the low pressure turbine 46. The
mid-turbine frame 57 further supports bearing systems 38 in the
turbine section 28. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via bearing systems 38 about the engine
central longitudinal axis A which is collinear with their
longitudinal axes.
[0037] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The mid-turbine
frame 57 includes airfoils 59 which are in the core airflow path C.
The turbines 46, 54 rotationally drive the respective low speed
spool 30 and high speed spool 32 in response to the expansion. It
will be appreciated that each of the positions of the fan section
22, compressor section 24, combustor section 26, turbine section
28, and fan drive gear system 48 may be varied. For example, gear
system 48 may be located aft of combustor section 26 or even aft of
turbine section 28, and fan section 22 may be positioned forward or
aft of the location of gear system 48.
[0038] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than about six (6), with an example embodiment being
greater than about ten (10), the geared architecture 48 is an
epicyclic gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3 and
the low pressure turbine 46 has a pressure ratio that is greater
than about five. In one disclosed embodiment, the engine 20 bypass
ratio is greater than about ten (10:1), the fan diameter is
significantly larger than that of the low pressure compressor 44,
and the low pressure turbine 46 has a pressure ratio that is
greater than about five (5:1). Low pressure turbine 46 pressure
ratio is pressure measured prior to inlet of low pressure turbine
46 as related to the pressure at the outlet of the low pressure
turbine 46 prior to an exhaust nozzle. The geared architecture 48
may be an epicycle gear train, such as a planetary gear system or
other gear system, with a gear reduction ratio of greater than
about 2.3:1. It should be understood, however, that the above
parameters are only exemplary of one embodiment of a geared
architecture engine and that the present invention is applicable to
other gas turbine engines including direct drive turbofans.
[0039] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft, with the engine at its best
fuel consumption--also known as "bucket cruise Thrust Specific Fuel
Consumption (`TSFC`)"--is the industry standard parameter of lbm of
fuel being burned divided by lbf of thrust the engine produces at
that minimum point. "Low fan pressure ratio" is the pressure ratio
across the fan blade alone, without a Fan Exit Guide Vane ("FEGV")
system. The low fan pressure ratio as disclosed herein according to
one non-limiting embodiment is less than about 1.45. "Low corrected
fan tip speed" is the actual fan tip speed in ft/sec divided by an
industry standard temperature correction of [(Tram .degree.
R)/(518.7.degree. R)].sup.0.5. The "Low corrected fan tip speed" as
disclosed herein according to one non-limiting embodiment is less
than about 1150 ft/second.
[0040] FIG. 2 shows a prior art airfoil 100 which may be
incorporated into an engine, such as engine 20 of FIG. 1 it its
turbine section. The airfoil 100 is illustrated having a suction
side 98 and a pressure side 99, and extending from a leading edge
97 to a trailing edge 93. The airfoil is illustrated as a blade
having a platform 11 at a radially inner end. The leading edge 97
is illustrated at the forward most end of the airfoil. A plurality
of showerhead cooling holes 102 are located along this leading edge
97. So-called gill cooling holes 106 are located spaced onto the
suction side 98 from the leading edge 97.
[0041] The showerhead cooling holes 102 may be defined as having a
longer dimension and a shorter dimension with the longer dimension
generally aligned with a radially outer direction. The gill cooling
holes also have a longer dimension and a shorter dimension,
however, the longer dimension in the gill holes 106 is generally
perpendicular to the radial dimension.
[0042] FIG. 4A illustrates the showerhead cooling holes 102 at the
leading edge 97. As shown, the gas flow path 120 wraps around the
leading edge and carries air onto the suction side with the prior
art showerhead cooling holes 102, the airflow 120, particularly at
the radially inner locations, would result in flow losses from the
air exiting the cooling holes 102.
[0043] As illustrated in FIG. 3, the leading edge 116 is spaced
from a trailing edge 115 and the suction side 98 is spaced from the
pressure side 101. However, contrary to the FIG. 2 airfoil, with
this high lift airfoil, there is an aerodynamic leading edge 118
that curves away from the mechanical leading edge 116 and onto the
pressure side 101. Thus, the aerodynamic leading edge has a portion
114 spaced relatively far from the mechanical leading edge 116 at
radially inner locations wherein the airfoil 101 merges into the
platform 111.
[0044] While the radially outer holes 113 may still be showerhead
cooling holes with a longer dimension generally aligned with the
radial dimension, there are transition cooling holes 115 which are
spaced intermediate a showerhead direction and a "gill" direction
at radially inner cooling holes 117.
[0045] As the airflow 120 extends from the pressure side around the
mechanical leading edge 116, the direction of the air exiting the
holes 130 and 132 will not result in the energy losses as would
occur if the prior art showerhead designs were utilized.
[0046] FIG. 3 shows a high lift blade 110. A high lift blade could
be defined as having an airfoil that maximizes the energy captured
from the working fluid. Working fluid flowing over the airfoil
surface exhibits a ratio of static pressure to total pressure in
proximity to the airfoil surface that is greater than approximately
0.9 across a substantial portion of the airfoil surface.
[0047] Refer to FIG. 4B, the radially outer showerhead holes 113
having a longer dimension d.sub.1 and a shorter dimension d.sub.2.
The longer dimension d.sub.1 is generally parallel to the radial
direction R. There is then a radially intermediate transition area
115 wherein the direction of the longer dimension d.sub.1 of a
cooling hole 130 is at an angle A with regard to the radial
dimension R. As shown at 132, the angle B may increase as the
location of the cooling holes extend radially inwardly through the
transition zone 115. Transition zone 115 includes a plurality of
cooling holes 130, 132 which have a longer dimension defining an
angle between 0 and 90.degree. relative to the radial
dimension.
[0048] As shown, there are a plurality of radially inner gill
cooling holes 117 which have their longer dimension d.sub.1
generally perpendicular to the radial direction R. The holes 113,
130, 132 and 117 are all non-circular with a longer dimension and a
shorter dimension.
[0049] With the disclosed positioning, the cooling air is more
efficiently utilized than in the prior art. While the location of
holes 113, 130, 132 and 117 may be at the leading edge, it is
envisioned that the holes need only be adjacent the leading edge.
The term "adjacent" as utilized with regard to this application is
defined as within 15% of the actual mechanical leading edge along a
surface length of either the pressure or suction side, as measured
from the leading edge toward the trailing edge.
[0050] Stated another way, an airfoil with a suction side 98 and
pressure side 101 extends from a leading edge 116 to a trailing
edge 115. A plurality of cooling holes 113, 117 are adjacent the
leading edge, with the cooling holes having a non-circular shape,
with a longer dimension and a smaller dimension. The airfoil
defines a radial direction R from a radially outer end to a
radially inner end. Radially outer cooling holes 113 spaced toward
the radially outer end have the longer dimension extending closer
to parallel to the radial direction, than radially inner cooling
holes 117, which are closer to the radially inner end. The radially
inner cooling holes 117 have the longer dimension extending to be
closer to perpendicular relative to the radial direction compared
to the radially outer cooling holes 113.
[0051] While this application has described the orientation of the
cooling holes with regard to a radial direction, in fact the
orientation of the holes is selected to more closely be orientated
with a main gas path flow direction at each location. However, in
the disclosed embodiment, the main gas path flow direction is
orientated so as to result in the hole orientation as described
above relative to the radial direction.
[0052] Although an embodiment of this invention has been disclosed,
a worker of ordinary skill in this art would recognize that certain
modifications would come within the scope of this invention. For
that reason, the following claims should be studied to determine
the true scope and content of this invention.
* * * * *